In gas turbines it is an aim to improve the efficiency of the gas turbine and in particular the efficiency of a gas turbine cycle, i.e. to increase the pressure ratio and the fired temperatures in a gas turbine. One way of increasing the efficiency of a given gas turbine cycle is to reduce the amount of cooling air used and thus allowing a longer expansion of high energy working fluid through the turbine resulting in an increase of the mechanical work that can be extracted from the turbine shaft.
A complex cooling system for a gas turbine with a high degree of tailoring in the distribution of cooling fluid over gas turbine components to be cooled is one way of reducing the total cooling air consumption. However, this method may give proper results for large scale components and geometries. When the size of the turbine components to be cooled decreases the relative size of e.g. the cooling apertures in comparison to the size of the components increase and the gain from sophisticated cooling air distribution techniques diminishes. The reason that the sizes of the cooling apertures cannot be further reduced are due to limitations in machining and manufacturing techniques of smaller turbine components and also in mitigating the risk that particles in the ambient air or from the internal component surfaces of the gas turbine will block narrow apertures and passages of such an internal cooling system with smaller apertures. As a consequence, a smaller gas turbine tends to use relatively more cooling air compared to larger gas turbines for the same gas turbine cycle and becomes therefore less efficient in respect of specific power and efficiency.
Hence, it is an aim to reduce the cooling fluid consumption. One way of reducing the cooling fluid consumption may be the use of serial cooling techniques where the same cooling fluid flow cools several parts of a component or of several different components.
DE 196 29 191 A1 discloses a device and a method for cooling a gas turbine with a ring-type combustion chamber. From a radially outer side a cooling fluid is fed into a ring channel. From the ring channel, the cooling fluid flows into a tube which is installed into a vane which is located at an exhaust section of the ring-type combustion chamber. A part of the cooling fluid flows from the ring channel along the radially outer wall of the combustion chamber and a further part of the cooling fluid which flows through the tube flows along a radially inner wall along the combustion chamber. The further part of the cooling fluid which has been flown through the tube may be guided to a blade which is located downstream and adjacent to the vane.
U.S. Pat. No. 5,953,919 A discloses a gas turbine comprising guide blades arranged between a combustion chamber and a turbine rotor. Guide blades are integrated in a respective associated combustion chamber of the gas turbine, wherein the guide blades and the associated combustion chamber wall are designed essentially in one piece and are constructed as a combustion chamber/guide blade unit. The unit is located at cool supporting structures of a gas turbine, wherein both together form cooling air passages which allow the combustion chamber to be cooled.
U.S. Pat. No. 5,320,485 A discloses a guide vane with a plurality of cooling circuits. The cooling circuits in the guide vane guides a cooling fluid through the interior of the vane while at the same time the pressure drop of the cooling fluid as it passes through the vane is minimized.
U.S. Pat. No. 5,592,820 A discloses a gas turbine diffuser for a turbo-jet engine, in which a plurality of diffuser inlets are arranged in a substantially circular array bounded by inner and outer diffuser walls and by a plurality of partitions extending generally radially from the axis of the annular diffuser.
U.S. Pat. No. 5,839,283 discloses mixing ducts for a gas turbine annular combustion chamber. In the gas turbine combustion chamber a row of pre-mixed burners are arranged in an annular form. A combustion air duct is designed as a diffuser which guides cooling fluid to each burner directly downstream of the compressor row.